Auxiliary power unit with electrically driven compressor

ABSTRACT

An auxiliary power unit for an aircraft includes a rotary intermittent internal combustion engine, a turbine having an inlet in fluid communication with an outlet of the engine, the turbine compounded with the engine, a compressor having an inlet in fluid communication with an environment of the aircraft and an outlet in fluid communication with the aircraft, the compressor rotatable independently of the turbine, an electric motor drivingly engaged to the compressor, and a transfer generator drivingly engaged to the engine, the transfer generator and the electric motor being electrically connected to allow power transfer therebetween. The compressor or an additional compressor may be in fluid communication with the inlet of the engine. A method of operating an auxiliary power unit of an aircraft is also discussed.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a divisional of U.S. application Ser. No.15/227,318, filed Aug. 3, 2016, which claims priority from U.S.application No. 62/202,283 filed Aug. 7, 2015, the entire contents ofwhich are incorporated by reference herein.

TECHNICAL FIELD

The application relates generally to compound engine assemblies and,more particularly, to such engine assemblies used as auxiliary powerunits in aircraft.

BACKGROUND OF THE ART

In auxiliary power units having a compressor and a turbine mechanicallyconnected to one another and/or to an engine core, the compressor,turbine and engine core are typically sized to be able to accommodatepeak loads. Accordingly, one or more components may not be operatingefficiently during operation at average loads. Moreover, a compromisemay need to be reached between the speed of the compressor and of theturbine, preventing one or both from operating at optimal speeds, whichmay compromise the aerodynamic efficiency of the auxiliary power unit.

Moreover, the compressor must typically be positioned in proximity ofthe component(s) to which they are mechanically connected, due to thegeometric constraints of shafts, gearing, etc. This may limit thepossible configurations for the auxiliary power unit, as well as theposition of its main inlet, with which the compressor is fluidlyconnected.

SUMMARY

In one aspect, there is provided an auxiliary power unit for anaircraft, comprising: an engine configured as a rotary intermittentinternal combustion engine; a turbine having an inlet in fluidcommunication with an outlet of the engine, the turbine compounded withthe engine; a compressor having an inlet in fluid communication with anenvironment of the aircraft and an outlet in fluid communication with ableed duct for providing bleed air to the aircraft, the compressorrotatable independently of the turbine; an electric motor drivinglyengaged to the compressor; and a transfer generator drivingly engaged tothe engine, the transfer generator and the electric motor beingelectrically connected to allow power transfer therebetween.

In another aspect, there is provided an auxiliary power unit for anaircraft, comprising: a plenum in fluid communication with anenvironment of the aircraft through a main inlet; an engine configuredas a rotary intermittent internal combustion engine; a turbine having aninlet in fluid communication with an outlet of the engine, the turbinecompounded with the engine; a first compressor having an inlet in fluidcommunication with the plenum and an outlet in fluid communication withan inlet of the engine; a second compressor having an inlet in fluidcommunication with the plenum and an outlet in fluid communication witha bleed duct for providing bleed air to the aircraft; an electric motordrivingly engaged to one of the first and second compressors, the one ofthe first and second compressors rotatable independently of the turbine;and a transfer generator drivingly engaged to the engine, the transfergenerator and the electric motor being electrically connected to allowpower transfer therebetween.

In a further aspect, there is provided a method of operating anauxiliary power unit of an aircraft, the method comprising: electricallydriving a compressor to provide compressed air to the aircraft;generating electrical power with a rotary intermittent internalcombustion engine; driving a turbine with an exhaust of the rotaryintermittent internal combustion engine; generating electrical powerwith the turbine; and transferring electrical power between thecompressor and the rotary intermittent internal combustion engine.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1a is a schematic view of an auxiliary power unit in accordancewith a particular embodiment;

FIG. 1b is a schematic view of an auxiliary power unit in accordancewith another particular embodiment;

FIG. 1c is a schematic view of an auxiliary power unit in accordancewith another particular embodiment;

FIG. 2 is a schematic cross-sectional view of a rotary engine which canbe used on the engine assembly of FIGS. 1a -1 c;

FIG. 3 is a schematic view of an auxiliary power unit in accordance withanother particular embodiment;

FIG. 4 is a schematic view of an auxiliary power unit in accordance withanother particular embodiment; and

FIG. 5 is a schematic view of an auxiliary power unit in accordance withanother particular embodiment.

DETAILED DESCRIPTION

Referring to FIG. 1a , an electric hybrid compound engine assembly 10used as an auxiliary power unit (APU) for airborne APU applications isgenerally shown. In a particular embodiment, the assembly 10 may beinstalled and removed as a single assembly like a conventional APU, withthe exception of a battery 8 thereof which may require a controlledtemperature environment. The battery 8 may be provided in a separatebattery compartment similar to that included in conventional APUinstallations, where the battery is typically used to power up the APUelectronic control and provide power for the starter. Alternately, theauxiliary power unit 10 may be used as a fixed or mobile ground powerunit.

The auxiliary power unit 10 includes an engine core 12′ including one ormore intermittent internal combustion engines 12 engaged to a commonshaft 16. In a particular embodiment, the intermittent internalcombustion engine(s) 12 is/are rotary internal combustion engine(s), forexample Wankel engine(s); it is however understood that other types ofinternal combustion engines may alternately be used. Referring to FIG.2, an example of a Wankel engine which may be used in the engine core12′ is shown. It is understood that the configuration of the engine(s)12, e.g. placement of ports, number and placement of seals, etc., mayvary from that of the embodiment shown.

The engine 12 comprises a housing 32 defining a rotor cavity having aprofile defining two lobes, which is preferably an epitrochoid. A rotor34 is received within the rotor cavity. The rotor defines threecircumferentially-spaced apex portions 36, and a generally triangularprofile with outwardly arched sides. The apex portions 36 are in sealingengagement with the inner surface of a peripheral wall 38 of the housing32 to form and separate three working chambers 40 of variable volumebetween the rotor 34 and the housing 32. The peripheral wall 38 extendsbetween two axially spaced apart end walls 54 to enclose the rotorcavity.

The rotor 34 is engaged to an eccentric portion 42 of an output shaft 16to perform orbital revolutions within the rotor cavity. The output shaft16 performs three rotations for each orbital revolution of the rotor 34.The geometrical axis 44 of the rotor 34 is offset from and parallel tothe axis 46 of the housing 32. During each orbital revolution, eachchamber 40 varies in volume and moves around the rotor cavity to undergothe four phases of intake, compression, expansion and exhaust.

An intake port 48 is provided through the peripheral wall 38 foradmitting compressed air into one of the working chambers 40. An exhaustport 50 is also provided through the peripheral wall 38 for discharge ofthe exhaust gases from the working chambers 40. Passages 52 for a sparkplug, glow plug or other ignition mechanism, as well as for one or morefuel injectors of a fuel injection system (not shown) are also providedthrough the peripheral wall 38. Alternately, the intake port 48, theexhaust port 50 and/or the passages 52 may be provided through the endor side wall 54 of the housing. A subchamber (not shown) may be providedin communication with the chambers 40, for pilot or pre injection offuel for combustion.

For efficient operation the working chambers 40 are sealed byspring-loaded peripheral or apex seals 56 extending from the rotor 34 toengage the inner surface of the peripheral wall 38, and spring-loadedface or gas seals 58 and end or corner seals 60 extending from the rotor34 to engage the inner surface of the end walls 54. The rotor 34 alsoincludes at least one spring-loaded oil seal ring 62 biased against theinner surface of the end wall 54 around the bearing for the rotor 34 onthe shaft eccentric portion 42.

The fuel injector(s) of the engine 12, which in a particular embodimentare common rail fuel injectors, communicate with a source of heavy fuel(e.g. diesel, kerosene (jet fuel), equivalent biofuel), and deliver theheavy fuel into the engine 12 such that the combustion chamber isstratified with a rich fuel-air mixture near the ignition source and aleaner mixture elsewhere.

Referring back to FIG. 1a , the auxiliary power unit 10 generallyincludes a supercharger compressor 20 compressing the air to feed theengine core 12′. Air from the environment of the aircraft enters aplenum 19 from the main inlet 14 of the auxiliary power unit 10. Thisplenum 19 feeds the compressor 20 so that an inlet of the compressor 20is in fluid communication with the environment of the aircraft. Thecompressor 20 optionally has variable inlet guide vanes 23 for flowcontrol and/or a variable diffuser 25. In a particular embodiment, thevariable diffuser 25 provides a low flow high pressure mode if thecompressor 20 is used to provide air to the aircraft when the enginecore 12′ is shut down.

An electric motor 64 (e.g. high speed electric motor) is drivinglyengaged to the compressor 20, for example by having a shaft of the motor64 directly connected to a shaft of the compressor 20, which is in turndirectly connected to the rotor(s) of the compressor 20. In a particularembodiment, the motor 64 is an alternating current constant speed drive;alternately, the motor 64 may be a variable speed drive. The type ofmotor is selected depending on the range of air flow and pressure outputrequired by the particular application for the auxiliary power unit 10.

The outlet of the compressor 20 is in fluid communication with the inletof the engine core 12′, in a particular embodiment through a heatexchanger 66: the compressor 20 thus delivers air through the heatexchanger 66 defining an intercooler and to the inlet of the engine core12′, for example to the intake port 48 of each rotary engine 12. In aparticular embodiment, the compressor 20 is located in proximity of theintercooler 66 for minimum ducting loss and weight. In the embodimentshown, the intercooler 66 is received in a cooling air duct 68 receivingair for example from a compartment of the auxiliary power unit 10, sothat cooling air may circulate through the intercooler 66 in heatexchange relationship with the compressed air fed to the engine core12′. Alternately, the intercooler 66 may be cooled through anintermediate fluid link to a main engine cooler 70 through which theused coolant from the engine core 12′ is circulated. The cooledcompressed air is delivered to the engine core 12′, for example at atemperature of 250° F. or less for an engine core 12′ including rotaryengine(s) 12.

The supercharger compressor 20 may also provide bleed air for theaircraft; in that case, air for the aircraft system is bled off beforethe intercooler 66, for example through a bleed air duct 72 as shown. Ina particular embodiment, a shut off valve (not shown) is providedupstream of the intercooler 66.

Alternately, for example when the flow-pressure requirements of theengine core 12′ and the aircraft cannot be efficiently reconciled to asingle compressor, an additional compressor 21 is provided to providethe bleed air to bleed air duct 72 for the aircraft, and the outlet ofthe compressor 20 communicates only with the inlet of the engine core12′ (i.e. the communication between the compressor 20 and the bleed airduct 72 is omitted). An excess air duct (now shown) connecting the bleedair duct 72 to the cooling air duct 68 through a valve may be providedto divert excess air to the cooling air duct when bleed air is notrequired by the aircraft, to prevent surge of the bleed compressor 21.The bleed compressor 21 may be driven by a separate motor 65 (as shownin dotted lines in FIG. 1a ) or by the same motor 64 as the corecompressor 20; alternately, one of the compressors 20, 21 can bemechanically coupled to the engine core 12′ running at nominallyconstant speed and the other electrically driven (as shown in FIG. 1b ),or one of the compressors 20, 21 can be coupled to the turbine sectionand the other electrically driven (as shown in FIG. 1c ).

In a particular embodiment, the cooling system includes the main engineliquid cooler 70, the intercooler 66 and an oil cooler 71. These aremounted close to the engine core 12′, for example in the cooling airduct 68 on a frame attached to the engine core 12′; the coolers 66, 70,71 may be mounted in series or in parallel. A fan 74 is located in thecooling air duct 68 downstream of the coolers 66, 70, 71 to drive(“pull”) airflow from the engine compartment through the cooling airduct 68 and the coolers 66, 70, 71 and into the tailpipe section. Inground operation there is a separate ventilation air inlet (not shown)to the compartment. In flight, when there is ram pressure, this aircomes from a side port (not shown) from the main aircraft inlet 14.Other configurations are also possible. In the embodiment shown, the fan74 is mechanically driven by the engine core 12′, for example through adirect engagement with the shaft 16 of the engine core 12′ such as torotate at a same speed.

The engine core 12′ is engaged to a transmission 28 which in turnsupports engine driven accessories (such as fuel and oil pumps, notshown). Although not shown, the fan 74 may also be driven through thetransmission 28. In a particular embodiment, an aircraft generator 76 isdirectly driven by the engine core 12′, for example including rotaryengine(s) 12 and with the engine shaft 16 rotating at 8000 rpm.Alternately, the aircraft generator 76 may be driven through step upgearing by the transmission 28, which may make the generator 76 morecompact.

The transmission 28 is also engaged to a transfer motor/generator 78(e.g. high speed motor/generator), the primary purpose of which is toeffect power transfer between the engine core 12′ and the rest of theassembly. The transfer motor/generator 78 may transfer power away fromthe engine core shaft 16 or temporarily to the engine core shaft 16 ifneeded. It can serve as a starter. Alternately, this power transferfunctionality may be integrated with the aircraft generator 76. However,in a particular embodiment separate motor/generators 76, 78 allow forimproved system segregation and failure tolerance.

The outlet of the engine core 12′ (e.g. exhaust port 50 of each rotaryengine 12) is in fluid communication with the inlet of a turbinesection, so that the exhaust from the engine core 12′ is fed to one ormore turbines 26, 22. One or more of the turbines 26, 22 is/areconfigured to compound power with the engine core 12′ (e.g. throughelectrical power transfer). In a particular embodiment, the turbines 26,22 are located as close to the engine core 12′ as possible to minimizehot ducting surface, pressure loss and weight. In a particularembodiment, the first stage turbine 26 has an outlet in fluidcommunication with an inlet of the second stage turbine 22, with theturbines 26, 22 having different reaction ratios from one another. Thedegree of reaction of a turbine can be determined using thetemperature-based reaction ratio (equation 1) or the pressure-basedreaction ratio (equation 2), which are typically close to one another invalue for a same turbine, and which characterize the turbine withrespect to “pure impulse” or “pure reaction” turbines:

$\begin{matrix}{{{Reaction}(T)} = \frac{( {t_{S\; 3} - t_{S5}} )}{( {t_{S\; 0} - t_{S\; 5}} )}} & (1) \\{{{Reaction}(P)} = \frac{( {P_{S3} - P_{S5}} )}{( {P_{S\; 0} - P_{S\; 5}} )}} & (2)\end{matrix}$where T is temperature and P is pressure, s refers to a static port, andthe numbers refers to the location the temperature or pressure ismeasured: 0 for the inlet of the turbine vane (stator), 3 for the inletof the turbine blade (rotor) and 5 for the exit of the turbine blade(rotor); and where a pure impulse turbine would have a ratio of 0 (0%)and a pure reaction turbine would have a ratio of 1 (100%).

In a particular embodiment, the first stage turbine 26 is configured totake benefit of the kinetic energy of the pulsating flow exiting thecore engine(s) 12 while stabilizing the flow and the second stageturbine 22 is configured to extract energy from the remaining pressurein the flow. Accordingly, in a particular embodiment the first stageturbine 26 has a lower reaction ratio (i.e. lower value) than that ofthe second stage turbine 22. In a particular embodiment, the first stageturbine 26 has a reaction ratio of 0.25 or lower (temperature orpressure based), and the second stage turbine 22 a reaction ratio higherthan 0.25 (temperature or pressure based). Other values are alsopossible.

One or more of the turbine stages 26, 22 powers a generator 80 (e.g.high speed generator). In the embodiment shown, both turbine stages 26,22 are drivingly engaged to the generator 80, for example by having ashaft of the generator 80 directly connected to a shaft of the turbines26, 22, which is in turn directly connected to the rotor(s) of theturbines 26, 22. Alternately, the generator 80 may be coupled to onlyone of the turbine stages 26, 22 while the other may be coupled to theengine core 12′ via the transmission or independently to a compressor20, 21. More than two turbine stages may also be provided.

Exhaust from the turbine stages 26, 22 is ducted to mix with the coolingsystem exhaust and out of the exhaust tailpipe 30.

In a particular embodiment, a power electronics module 82 is locatedclose to the cooling air inlet 69. The power electronics module 82provides an electrical connection between the transfer motor/generator78, the compressor motor(s) 64, 65, the turbine generator 80 and thebattery 8. In a particular embodiment, the power electronics module 82contains an alternating current motor drive for the compressor motor(s)64, 65, rectifier and regulator for the turbine generator 80,bi-directional drive/regulation for the transfer motor/generator 78 andcharge current regulator for the battery 8; alternately, the chargecurrent regulator may be located within the battery assembly. The powerelectronics module 82 accepts inputs from an APU electronic control 84,which in turn responds to the aircraft input and feedback from systemsensors.

In a particular embodiment, the storage battery 8 has high energydensity, and is for example a Li-Polymer multi cell 270 V direct currentbattery. In a particular embodiment, this battery 8 is capable ofoutputs up to 50 or 60 KW over very short periods, and an energy densityaround 0.75 KW/Kg. Other values are also possible.

FIG. 1b shows an electric hybrid compound engine assembly auxiliarypower unit 110 in accordance with another embodiment, where elementssimilar to that of the embodiments of FIG. 1a are identified with thesame reference numerals and will not be further described herein. Inthis embodiment the bleed compressor 121 is driven by an electric motor165 while the core compressor 120 is mechanically coupled to the enginecore 12′, either directly or through the transmission 28.

FIG. 1c shows an electric hybrid compound engine assembly auxiliarypower unit 110′ in accordance with another embodiment, where elementssimilar to that of the embodiments of FIG. 1a are identified with thesame reference numerals and will not be further described herein. Inthis embodiment the bleed compressor 121′ is driven by an electric motor165 while the core compressor 120′ is mechanically coupled to theturbines 22, 26, for example by being coupled to the turbine generator80 as shown.

FIG. 3 shows an electric hybrid compound engine assembly auxiliary powerunit 210 in accordance with another embodiment, where elements similarto that of the embodiments of FIGS. 1a-1b are identified with the samereference numerals and will not be further described herein. Althoughnot shown, an additional compressor 21 could be provided for aircraftbleed air, with the compressors 20, 21 configured for example as setforth in FIG. 1a, 1b or 1 c.

In this embodiment, the cooling fan 274 is electrically powered by a fanmotor 286, rather than by a mechanical drive connected to the enginecore 12′. In this configuration the power electronics module 82 providesan additional output for the fan motor 286, so that the fan motor 286 iselectrically connected to the transfer motor/generator 78, thecompressor motor(s) 64, the turbine generator 80 and the battery. Thecomplexity of this output may vary. It could be a simple direct currentlink with the motor control electronics on the fan motor 286, it couldbe an alternating current link or it could be bi-directional; in aparticular embodiment, a bi-directional link allows for harvesting ofpower from the wind milling of the fan 274 when the door of the inlet 69is open and the auxiliary power unit 210, 310, 410 is non-operatingduring flight.

This embodiment can provide some advantages in the sizing and operationof the fan 274: it may be turned off for example to save power until thecoolant reaches a predetermined temperature. Its operating speed may bereduced on cold days when less cooling effort is required. Its designspeed may be optimized for packaging volume.

FIG. 4 shows an electric hybrid compound engine assembly auxiliary powerunit 310 in accordance with another embodiment, where elements similarto that of the embodiment of FIG. 3 are identified with the samereference numerals and will not be further described herein. Althoughnot shown, an additional compressor 21 could be provided for aircraftbleed air, with the compressors 20, 21 configured for example as setforth in FIG. 1a, 1b or 1 c. Although the fan 274 is shown as beingdriven by the fan motor 286, it could alternately be driven by theengine core 12′, for example as set forth in FIG. 1 a.

In this embodiment, the turbine generator and compressor motor arereplaced by a common electrical machine 388 (e.g. high speedmotor/generator). The turbines 26, 22 are mechanically linked to thecompressor 20 by a turbo-machine shaft 324, and the motor/generator 388is engaged to the turbo-machine shaft 324. The motor/generator 388 actsas a motor or a generator depending on whether it is required to add orsubtract torque. In a particular embodiment, attaching the turbines 26,22 to the compressor 20 mechanically reduces the power transferrequirements of the system in normal operation, which is a considerationsince the motor/generator 388 and power electronics module 82 createheat in relation to the electric current being handled.

In a particular embodiment, the turbines 26, 22 are coupled to theturbo-machine shaft 324 by an over-running clutch 390 such that thecompressor 20 can be accelerated to a higher speed than the free runningspeed of the turbines 26, 22; the compressor 20 may thus selectively berotatable independently of the turbines 26, 22. In a particularembodiment, the clutch 390 allows for boosting the supercharge intransient conditions from low power of the engine core 12′ (to reduce oreliminate turbo lag) and for motoring the compressor 20 without creatingturbine drag when the engine core 12′ is shut down. In a particularembodiment, the clutch 390 provides some protection to the turbines 26,22 to reduce the risk of or prevent the turbines 26, 22 from exceedingthe set speed of the motor/generator 388, by allowing themotor/generator 388 and the compressor 20 to act as a brake on theturbines 26, 22 through the over-running clutch 390.

Alternately, the over-running clutch 390 may be replaced by any type ofclutch having positive engagement and disengagement configurationssuitable for use at the rotational speeds of the compressor 20 andturbines 26, 22. In a particular embodiment steps are taken to limit (orstructurally tolerate) the maximum speed that can be achieved byde-coupled turbines. In an alternate embodiment, the turbines 26, 22 andcompressors 20 are connected through an intermediate gear system withthe clutch (e.g. overrunning clutch) separating the electricalmachine/compressor portion of the drive from the turbine portion of thedrive; the clutch may be provided at a convenient intermediate gearstage having a speed compatible with the over-running clutch.

FIG. 5 shows an electric hybrid compound engine assembly auxiliary powerunit 410 in accordance with another embodiment, where elements similarto that of the embodiment of FIG. 4 are identified with the samereference numerals and will not be further described herein. Althoughnot shown, an additional compressor 21 could be provided for aircraftbleed air, with the compressors 20, 21 configured for example as setforth in FIG. 1a, 1b or 1 c. Although the fan 274 is shown as beingdriven by the fan motor 286, it could alternately be driven by theengine core 12′, for example as set forth in FIG. 1 a.

In this embodiment, the turbines 22, 426 are spilt. The first stage(e.g. lower reaction ratio) turbine 426 is more efficient at anintermediate speed and is compounded to the engine core 12′ by beingmechanically engaged therewith, for example through the transmission 28.The second stage (e.g. higher reaction ratio) turbine 22 is coupled tothe compressor 20 through a free turbo-machine shaft 424 in a similarmanner to a turbocharger. The motor/generator 388 on the turbo-machineshaft 424 allows for transferring power to the compressor 20 orextracting more load from the second stage turbine 22. The second stageturbine 22 is optionally connected to the turbo-machine shaft 424through a clutch 490 (as shown) so as to allow for motoring of thecompressor 20 without creating turbine drag when the engine core 12′ isshut down.

In use and in a particular embodiment, the auxiliary power unit 10, 110,110′, 210, 310, 410 is operated by electrically driving the compressor20, 21 to provide compressed air to the aircraft, generating electricalpower with the engine core 12′, driving the turbines 22, 26, 426 with anexhaust of the engine core, generating electrical power with theturbine(s) 22, 26, and transferring electrical power between thecompressor 20, 21 and the engine core 12′. The method may furtherinclude electrically driving the separate core compressor 20 to providecompressed air to the engine core 12′, and transferring electrical powerbetween the compressors 20, 21 and the engine core 12′.

In a particular embodiment, the use of electrical machines 64, 65, 78,80, 165, 388 on the turbines 22, 26, compressor(s) 20, 21 and enginecore 12′ and management of the energy transfer between the electricalmachines 64, 65, 78, 80, 165, 388 allows for the system to adapt tochanges in operating conditions. The storage of energy in the battery 8advantageously provides for the energy from the battery 8 to be used forpeak load topping (for example to support main engine start (MES) ormaximum demand on the aircraft environmental control system (ECS), whichrepresent the highest loads of an auxiliary power unit running on theground). Similarly the battery 8 can be used when the auxiliary powerunit 10, 110, 110′, 210, 310, 410 is used in flight, for example at thetop of descent at very high altitude (e.g. above 35,000 feet) whichdemands high load at low atmospheric pressure for a short period oftime. Consequently the loads can be averaged and the size of the enginecore 12′ can be optimized for an average duty cycle load includingbattery recharge, which may allow for the engine core 12′ to be smallerand operated for a larger percentage of the duty cycle at optimalefficiency conditions. This may allow a reduction of average fuelconsumption, emission and/or noise levels by having the engine core 12′operate at preferred conditions for a greater portion of the cycle.

In a particular embodiment, load transients may be improved by managingthe electric power transfer system between the engine core 12′ and thecompressor 20.

In a particular embodiment, the electrically driven compressor(s) 20, 21and electrically compounded turbine(s) 22, 26 can be run at optimalspeeds and therefore be smaller and/or more aerodynamically efficientthan compressor(s) and turbine(s) mechanically coupled to the rotarycore, where a compromise may be necessary.

In a particular embodiment, power can be transferred between the enginecore 12′ and the compressor 20, 21 to optimize the operating point ofeach. With variable speed operation of the compressor 20, 21 a widerange of flow and pressure may be obtained efficiently without recourseto inefficient highly closed inlet guide vane settings for lower flowand pressure ratio conditions.

In a particular embodiment and as discussed above, the compressor 20, 21can be individually driven electrically without requiring start-up ofthe engine core 12′. Accordingly, it may be possible to offer a shortterm pneumatic power capability to the aircraft (e.g. perform pneumaticmain engine starting and/or provide some cabin conditioning for a shortperiod) using the compressor 20, 21 driven only on battery power, i.e.without starting the auxiliary power unit 10, 110, 110′, 210, 310, 410.This may be advantageous for redundancy (e.g. for main engine start) andin situations where minimum noise and emissions are critical (e.g. nearthe gate area).

In particular embodiments where the compressor 20, 21 is driven by adedicated electrical motor 64, 65, 165, the compressor 20, 21 and inlet14 can be positioned as convenient for the benefit of the overallinstallation package, without worrying about geometric constraints ofshafts and gearing to the engine core shaft 16. Shaft lengths can belimited (for example) by dynamic concerns and gear center distances maybe limited by pitch line speeds; angles of elements like bevel drivesmay also be constrained. By contrast, transmitting the load electricallyis done by wires which may be routed conveniently as required, providingmore flexibility in positioning of the driven elements.

In particular embodiments, the electrically driven cooling fan 274allows the cooling fan 274 to be used as a ram air turbine in flight ifthe APU inlet door is opened while the auxiliary power unit 210, 310,410 is not operating. Electricity generated by the electrical machine286 engaged to the fan 274 and/or the power of the battery 8 can be usedin emergencies to eliminate the need for a separate ram air turbine usedfor last chance emergency power on many larger commercial aircraft; theelectricity generated by the fan 274 may be used to supplant starterpower in flight and/or achieve windmill starting of the engine core 12′without battery assist. Additionally the APU battery 8 may be used as asignificant source of aircraft emergency electrical power. In the eventimproved segregation is required for the emergency wind-millingfunction, a second set of windings may be included in the fan motor 286,dedicated to the emergency power function.

In a particular embodiment, the fan 74 mechanically driven by the enginecore 12′ as per FIGS. 1a-1b may also allow for windmill starting.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed without departing from the scope of the invention disclosed.Modifications which fall within the scope of the present invention willbe apparent to those skilled in the art, in light of a review of thisdisclosure, and such modifications are intended to fall within theappended claims.

The invention claimed is:
 1. A method of operating an auxiliary powerunit of an aircraft, the method comprising: electrically driving acompressor to provide compressed air to the aircraft; generatingelectrical power with a rotary intermittent internal combustion engine;driving a turbine with an exhaust of the rotary intermittent internalcombustion engine; generating electrical power with the turbine; andtransferring electrical power between the compressor and the rotaryintermittent internal combustion engine.
 2. The method as defined inclaim 1, wherein the compressor is a first compressor, the methodfurther comprising: electrically driving a second compressor to providecompressed air to an inlet of the rotary intermittent internalcombustion engine; and transferring electrical power between the firstcompressor, the second compressor and the rotary intermittent internalcombustion engine.
 3. The method as defined in claim 1, furthercomprising electrically driving a cooling fan with the electrical powergenerated with the rotary intermittent internal combustion engine orwith the electrical power generated with the turbine, the cooling fan influid communication with coolers providing heat exchange relationshipbetween a liquid coolant of the rotary intermittent internal combustionengine and ambient air.
 4. The method as defined in claim 1, furthercomprising transferring the electrical power generated with both of therotary intermittent internal combustion engine and with the turbine to apower electronic module.
 5. The method as defined in claim 4, furthercomprising selectively transferring the electrical power from the powerelectronic module to a first electric motor in driving engagement withthe compressor, to a second electric motor in driving engagement with acooling fan, to a battery, and/or to an accessory.
 6. The method asdefined in claim 1, wherein electrically driving the compressor includesdriving an electric motor directly drivingly engaged to a shaft of thecompressor.
 7. The method as defined in claim 2, wherein electricallydriving the second compressor includes driving the compressor and thesecond compressor with a common electric motor in driving engagementwith shafts of the compressor and the second compressor.
 8. The methodas defined in claim 2, wherein electrically driving the secondcompressor includes driving each of the compressor and the secondcompressor with a respective one of two electric motors.
 9. The methodas defined in claim 1, further comprising feeding compressed air to therotary intermittent internal combustion engine.
 10. The method asdefined in claim 1, wherein the turbine is a first stage turbine, themethod further comprising driving a second stage turbine having an inletin fluid communication with an outlet of the first stage turbine. 11.The method as defined in claim 10, wherein generating electrical powerwith the turbine includes driving a generator with the second stageturbine.
 12. The method as defined in claim 11, wherein driving theturbine with the exhaust of the rotary intermittent internal combustionengine includes driving a shaft of the engine with the first stageturbine.
 13. The method as defined in claim 1, wherein generating theelectrical power with the turbine includes driving a generator with theturbine through a clutch.
 14. The method as defined in claim 2, furthercomprising transferring heat from the compressed air exiting the secondcompressor to a flow of cooling air before feeding the compressed air tothe rotary intermittent internal combustion engine.
 15. The method asdefined in claim 14, further comprising generating the flow of coolingair by driving a cooling fan.
 16. The method as defined in claim 15,wherein driving the cooling fan includes driving the cooling fan with anelectric motor.